Thermal management system

ABSTRACT

A gas turbine engine includes a turbomachine and a thermal management system. The thermal management system includes a heat source heat exchanger configured to collect heat from the turbomachine during operation; a heat sink heat exchanger; and a thermal transport bus having a heat exchange fluid configured to flow therethrough at a pressure within an operational pressure range. The thermal management system defines an operational temperature range for the heat exchange fluid, the operational temperature range having a lower temperature limit less than about zero degrees Fahrenheit at a pressure within the operational pressure range and an upper temperature limit of at least about 1000 degrees Fahrenheit at a pressure within the operational pressure range.

FIELD

The present subject matter relates generally to a thermal managementsystem, and more specifically to a thermal management system for a gasturbine engine.

BACKGROUND

A gas turbine engine typically includes a fan and a turbomachine. Theturbomachine generally includes an inlet, one or more compressors, acombustor, and at least one turbine. The compressors compress air whichis channeled to the combustor where it is mixed with fuel. The mixtureis then ignited for generating hot combustion gases. The combustiongases are channeled to the turbine(s) which extracts energy from thecombustion gases for powering the compressor(s), as well as forproducing useful work to propel an aircraft in flight and/or to power aload, such as an electrical generator.

In at least certain embodiments, the turbomachine and fan are at leastpartially surrounded by an outer nacelle. With such embodiments, theouter nacelle defines a bypass airflow passage with the turbomachine.Additionally, the turbomachine is supported relative to the outernacelle by one or more outlet guide vanes/struts.

During operation of the gas turbine engine, various systems may generatea relatively large amount of heat. Thermal management systems of the gasturbine engine may collect heat from one or more of these systems tomaintain a temperature of such systems within an acceptable operatingrange. The thermal management systems may reject such heat through oneor more heat exchangers.

Typically, these thermal management systems are designed to operate witha designated subset of the systems within the gas turbine engine basedon such systems' operating temperatures and thermal management needs.However, such a configuration may lead to, e.g., an increased complexityand decreased efficiency.

Accordingly, the inventors of the present disclosure have found that athermal management system capable of operating over a wider temperaturerange would be useful to facilitate accommodation of the various systemswithin the gas turbine engine.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a gas turbineengine is provided. The gas turbine engine includes a turbomachine; anda thermal management system. The thermal management system includes afirst heat exchanger configured to collect heat from the turbomachineduring operation; a second heat exchanger; and a thermal transport bushaving a heat exchange fluid configured to flow therethrough at apressure within an operational pressure range, the first heat exchangerand second heat exchanger each fluidly coupled to the thermal transportbus such that the first heat exchanger is operable to transfer heat tothe heat exchange fluid during at least certain operations and thesecond heat exchanger is operable to transfer heat from the heatexchange fluid during at least certain operations, the thermalmanagement system defining an operational temperature range for the heatexchange fluid, the operational temperature range having a lowertemperature limit less than about zero degrees Fahrenheit at a pressurewithin the operational pressure range and an upper temperature limit ofat least about 1000 degrees Fahrenheit at a pressure within theoperational pressure range.

In certain exemplary embodiments the first heat exchanger is a heatrecovery heat exchanger.

For example, in certain exemplary embodiments the turbomachine includesa turbine section and an exhaust section, the turbine section andexhaust section together defining at least in part a core air flowpath,and wherein the heat recovery heat exchanger is a waste heat recoveryheat exchanger positioned to be in thermal communication with the coreair flowpath within or downstream of the turbine section, the exhaustsection, or both.

For example, in certain exemplary embodiments the turbomachine includesa fuel delivery system, and wherein the second heat exchanger is a fuelheat exchanger thermally coupled to the fuel delivery system.

In certain exemplary embodiments the turbomachine includes a cooledcooling air system, and wherein the first heat exchanger is thermallycoupled to the cooling air system.

In certain exemplary embodiments the lower temperature limit of theoperational temperature range at a pressure within the operationalpressure range is between about −100 degrees Fahrenheit and about −5degrees Fahrenheit.

For example, in certain exemplary embodiments the upper temperaturelimit of the operational temperature range at a pressure within theoperational pressure range is between about 1000 degrees Fahrenheit andabout 1800 degrees Fahrenheit.

In certain exemplary embodiments heat exchange fluid is at least one ofa liquid metal alloy, a molten salt, a silicone oil, an ionic fluid, apressurized gas, or a supercritical gas.

For example, in certain exemplary embodiments the operational pressurerange is greater than zero pounds per square inch and less than about500 pounds per square inch.

For example, in certain exemplary embodiments heat exchange fluid is aeutectic metal alloy including gallium, indium, and tin.

In certain exemplary embodiments the heat exchange fluid is apressurized gas.

For example, in certain exemplary embodiments the heat exchange fluid isa supercritical gas, wherein the supercritical gas defines criticalpoint pressure, wherein the operational pressure range is greater thanthe critical point pressure of the supercritical gas and up to about8000 pounds per square inch.

For example, in certain exemplary embodiments the heat exchange fluid isa supercritical gas, and wherein the operational pressure range isgreater than about 1000 pounds per square inch and less than about 8000pounds per square inch.

In certain exemplary embodiments, the gas turbine engine furtherincludes an outer nacelle at least partially surrounding theturbomachine; and an outlet guide vane extending between the outernacelle and the turbomachine, wherein the second heat exchanger isintegrated into, or coupled to, the outlet guide vane.

In certain exemplary embodiments the thermal management system furtherincludes a heater thermally coupled to the thermal transfer bus forheating the thermal transfer fluid within the thermal transfer bus.

In certain exemplary embodiments the first heat exchanger is a firstheat source heat exchanger, wherein the thermal management systemfurther includes a second heat source heat exchanger, wherein the firstheat source heat exchanger is a waste heat recovery heat exchanger, andwherein the second heat source heat exchanger is a cooling air systemheat exchanger.

In certain exemplary embodiments the second heat exchanger is a fuelheat exchanger, and wherein the thermal management system furthercomprises a bypass airflow passage heat exchanger.

In another exemplary embodiment of the present disclosure, gas turbineengine is provided. The gas turbine engine includes a compressorsection, a combustion section, a turbine section, and an exhaust sectionarranged in serial flow order and together defining a core air flowpath.The gas turbine engine also includes a fuel delivery system forproviding a flow of fuel to the combustion section. The gas turbineengine also includes a thermal management system. The thermal managementsystem includes a first heat exchanger positioned to be in thermalcommunication with a flow through the core air flowpath within ordownstream of the turbine section, the exhaust section, or both; asecond heat exchanger in thermal communication with the fuel deliverysystem for transferring heat to the flow of fuel provided to thecombustion section; and a thermal transport bus having a heat exchangefluid flowing therethrough, the first heat exchanger and secondexchanger each fluidly coupled to the thermal transport bus.

In certain exemplary embodiments the thermal management system definesan operational temperature range for the heat exchange fluid, theoperational temperature range having a lower temperature limit less thanabout zero degrees Fahrenheit at a pressure within the operationalpressure range and an upper temperature limit of at least about 1000degrees Fahrenheit at a pressure within the operational pressure range.

For example, in certain exemplary embodiments the heat exchange fluid isat least one of a molten salt, a silicone oil, an ionic fluid, a liquidmetal alloy, or a supercritical gas.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic, cross-sectional view of an exemplary gas turbineengine according to various embodiments of the present subject matter.

FIG. 2 is a simplified schematic view of a thermal management system inaccordance with an exemplary embodiment of the present disclosure.

FIG. 3 is a schematic, cross-sectional view of a section of a gasturbine engine including a thermal management system in accordance withan exemplary embodiment of the present disclosure.

FIG. 4 is a schematic view of a heater in accordance with an exemplaryembodiment of the present disclosure.

FIG. 5 is a schematic view of a heater in accordance with anotherexemplary embodiment of the present disclosure.

FIG. 6 is a schematic view of a heater in accordance with yet anotherexemplary embodiment of the present disclosure.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 10percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematic,cross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan engine 10includes a fan section 14 and a turbomachine 16 disposed downstream fromthe fan section 14.

The exemplary turbomachine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. The compressorsection, combustion section 26, turbine section, and exhaust nozzlesection 32 together define at least in part a core air flowpath 37through the turbomachine 16. A high pressure (HP) shaft or spool 34drivingly connects the HP turbine 28 to the HP compressor 24. A lowpressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 tothe LP compressor 22.

For the embodiment depicted, the fan section 14 includes a variablepitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 ina spaced apart manner. As depicted, the fan blades 40 extend outwardlyfrom disk 42 generally along the radial direction R. Each fan blade 40is rotatable relative to the disk 42 about a pitch axis P by virtue ofthe fan blades 40 being operatively coupled to a suitable actuationmember 44 configured to collectively vary the pitch of the fan blades 40in unison. The fan blades 40, disk 42, and actuation member 44 aretogether rotatable about the longitudinal axis 12 by LP shaft 36 acrossa power gear box 46. The power gear box 46 includes a plurality of gearsfor stepping down the rotational speed of the LP shaft 36 to a moreefficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, the disk 42 iscovered by rotatable front hub 48 aerodynamically contoured to promotean airflow through the plurality of fan blades 40. Additionally, theexemplary fan section 14 includes an annular fan casing or outer nacelle50 that circumferentially surrounds the fan 38 and/or at least a portionof the turbomachine 16. The nacelle 50 is supported relative to theturbomachine 16 by a plurality of circumferentially-spaced outlet guidevanes 52. Moreover, the nacelle 50 extends over an outer portion of theturbomachine 16 so as to define a bypass airflow passage 56therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersthe turbofan 10 through an associated inlet 60 of the nacelle 50 and/orfan section 14. As the volume of air 58 passes across the fan blades 40,a first portion of the air 58 as indicated by arrows 62 is directed orrouted into the bypass airflow passage 56 and a second portion of theair 58 as indicated by arrow 64 is directed or routed into the LPcompressor 22. The ratio between the first portion of air 62 and thesecond portion of air 64 is commonly known as a bypass ratio. As stated,for the embodiment shown, the turbofan engine 10 is a high bypassturbofan engine 10. Accordingly, for the embodiment depicted, the bypassratio defined by the turbofan engine 10 is greater than about 6:1 and upto about 30:1.

The pressure of the second portion of air 64 is then increased as it isrouted through the high pressure (HP) compressor 24 and into thecombustion section 26, where it is mixed with fuel and burned to providecombustion gases 66. Subsequently, the combustion gases 66 are routedthrough the HP turbine 28 and the LP turbine 30, where a portion ofthermal and/or kinetic energy from the combustion gases 66 is extracted.

The combustion gases 66 are then routed through the jet exhaust nozzlesection 32 of the turbomachine 16 to provide propulsive thrust.Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.

Moreover, as is depicted schematically, the exemplary turbofan engine 10further includes various accessory systems to aid in the operation ofthe turbofan engine 10 and/or an aircraft including the turbofan engine10 (see, e.g., FIG. 3). For example, the exemplary turbofan engine 10further includes a main lubrication system 78 configured to provide alubricant to, e.g., various bearings and gear meshes in the compressorsection (including the LP compressor 22 and HP compressor 24), theturbine section (including the HP turbine 28 and the LP turbine 30), theHP spool 34, the LP spool 36, and the power gear box 46. The lubricantprovided by the main lubrication system 78 may increase the useful lifeof such components and may remove a certain amount of heat from suchcomponents. Additionally, the turbofan engine 10 includes a cooling air(CCA) system 80 (sometimes also referred to as a “compressor cooling airsystem” or “cooled cooling air system”) for providing air from one orboth of the HP compressor 24 or LP compressor 22 to one or both of theHP turbine 28 or LP turbine 30. Moreover, the exemplary turbofan engine10 includes an active thermal clearance control (ACC) system 82 forcooling a casing of the turbine section to maintain a clearance betweenthe various turbine rotor blades and the turbine casing within a desiredrange throughout various engine operating conditions. Furthermore, theexemplary turbofan engine 10 includes a generator lubrication system 84for providing lubrication to an electronic generator, as well ascooling/heat removal for the electronic generator. The electronicgenerator may provide electrical power to, e.g., a startup electricmotor for the turbofan engine 10 and/or various other electroniccomponents of the turbofan engine 10 and/or an aircraft including theturbofan engine 10.

As is also depicted schematically, the exemplary turbofan engine 10depicted drives or enables various other accessory systems, e.g., for anaircraft (not shown) including the exemplary turbofan engine 10. Forexample, the exemplary turbofan engine 10 provides compressed air fromthe compressor section to an environmental control system (ECS) 86. TheECS 86 may provide an air supply to a cabin of the aircraft forpressurization and thermal control. Additionally, air may be providedfrom the exemplary turbofan engine 10 to an electronics cooling system88 for maintaining a temperature of certain electronic components of theturbofan engine 10 and/or aircraft within a desired range.

Prior turbofan engines 10 and/or aircrafts included individual heatexchangers for each of these accessory systems to remove heat from airand/or lubrication in such systems. However, aspects of the presentdisclosure may include a thermal management system 100 (see FIGS. 2 and3) for transferring heat from some or all of such accessory systems tomore efficiently remove such heat and/or utilize such heat.

It should be appreciated, however, that the exemplary turbofan engine 10depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, aspects of the present disclosure mayadditionally, or alternatively, be applied to any other suitable gasturbine engine. For example, in other exemplary embodiments, theturbofan engine 10 may instead be any other suitable aeronautical gasturbine engine, such as a turbojet engine, turboshaft engine, turbopropengine, etc. Additionally, in still other exemplary embodiments, theexemplary turbofan engine 10 may include or be operably connected to anyother suitable accessory systems and may be configured in any othersuitable manner. Additionally, or alternatively, the exemplary turbofanengine 10 may not include or be operably connected to one or more of theaccessory systems discussed above.

Referring now to FIG. 2, a schematic, flow diagram is provided of athermal management system 100 in accordance with an exemplary embodimentof the present disclosure for incorporation at least partially into agas turbine engine, such as the exemplary turbofan engine 10 of FIG. 1.

As shown, the thermal management system 100 generally includes a thermaltransport bus 102. The thermal transport bus 102 includes anintermediary heat exchange fluid flowing therethrough and may be formedof one or more suitable fluid conduits. The heat exchange fluid may havea high temperature operating range, as will be explained in more detailbelow. A pump 104 is provided in fluid communication with the heatexchange fluid in the thermal transport bus 102 for generating a flow ofthe heat exchange fluid in/through the thermal transport bus 102. Asviewed in FIG. 2, the pump 104 may generate a flow of the heat exchangefluid generally in a clockwise direction through the thermal transportbus 102. The pump 104 may be a rotary pump including an impeller, oralternatively may be any other suitable fluid pump. Additionally, thepump 104 may be powered by an electric motor, or alternatively may be inmechanical communication with and powered by, e.g., the HP shaft 34 orthe LP shaft 36 of the turbofan engine 10. In still other embodiments,the pump 104 may be powered by an auxiliary turbine, which in turn maybe powered by bleed air from a compressor section of a gas turbineengine within which the system 100 is incorporated.

Moreover, the exemplary thermal management system 100 includes one ormore heat source heat exchangers 106 in thermal communication with thethermal transport bus 102, or rather, in fluid communication with theheat exchange fluid within the thermal transport bus 102. Specifically,the thermal management system 100 depicted includes a plurality of heatsource heat exchangers 106. The plurality of heat source heat exchangers106 are each configured to transfer heat from one or more of theaccessory systems of the turbofan engine 10 (or operable with theturbofan engine 10) to the heat exchange fluid in the thermal transportbus 102. For example, in certain exemplary embodiments, the plurality ofheat source heat exchangers 106 may include one or more of: a heatrecovery heat exchanger, such as a waste heat recovery heat exchanger,positioned in, e.g., the turbine section or exhaust section forrecovering heat from an airflow therethrough; a CCA system heat sourceexchanger for transferring heat from a CCA system (such as CCA system80); a main lubrication system heat exchanger for transferring heat froma main lubrication system (such as main lubrication system 78); an ACCsystem heat source exchanger for transferring heat from an ACC system(such as ACC system 82); a generator lubrication system heat sourceexchanger for transferring heat from a generator lubrication system(such as generator lubrication system 84); an ECS heat exchanger fortransferring heat from an ECS (such as ECS 86); an electronics coolingsystem heat exchanger for transferring heat from an electronics coolingsystem (such as electronics cooling system 88); a vapor compressionsystem heat exchanger; an air cycle system heat exchanger; and anauxiliary system(s) heat source exchanger. By way of example, theauxiliary system(s) heat source exchanger may be configured to transferheat from one or more of a radar system, a defense system, passengerentertainment systems, etc. Accordingly, a thermal management system 100in accordance with an exemplary embodiment of FIG. 2 may transfer heatfrom a variety of independent systems to the heat exchange fluid in thethermal transport bus 102 for removal.

For the embodiment depicted, there are three heat source heat exchangers106, the three heat source heat exchangers 106 each arranged in seriesflow along the thermal transport bus 102. However, in other exemplaryembodiments, any other suitable number of heat source heat exchangers106 may be included and one or more of the heat source heat exchangers106 may be arranged in parallel flow along the thermal transport bus102. For example, in other embodiments, there may be a single heatsource exchanger 106 in thermal communication with the heat exchangefluid in the thermal transport bus, or alternatively, there may be atleast two heat source heat exchangers 106, at least four heat sourceheat exchangers 106, at least five heat source heat exchangers 106, orat least six heat source heat exchangers 106 in thermal communicationwith heat exchange fluid in the thermal transport bus 102.

Additionally, the exemplary thermal management system 100 of FIG. 2further includes one or more heat sink exchangers 108 in thermalcommunication with the thermal transport bus 102, or rather in fluidcommunication with the heat exchange fluid in the thermal transport bus102. The one or more heat sink exchangers 108 are located downstream ofthe plurality of heat source exchangers 106 and are configured fortransferring heat from the heat exchange fluid in the thermal transportbus 102, e.g., to atmosphere, to fuel, to a fan stream, etc. Forexample, in certain embodiments the one or more heat sink exchangers 108may include at least one of a fuel heat exchanger, a fan stream heatexchanger, a RAM heat exchanger, a bleed air heat exchanger, an engineintercooler, or a cold air output of an air cycle system. The fuel heatexchanger may be a “fluid to heat exchange fluid” heat exchanger whereinheat from the heat exchange fluid is transferred to a stream of liquidfuel for the turbofan engine 10 (by, e.g., a fuel delivery system).Moreover, the fan stream heat exchanger may generally be an “air to heatexchange fluid” heat exchanger which flows, e.g., bypass air from abypass airflow passage over heat exchange fluid to remove heat from theheat exchange fluid. Additionally, the RAM heat exchanger may beconfigured as an “air to heat exchange fluid” heat exchanger integratedinto one or both of the turbofan engine 10 or an aircraft including theturbofan engine 10. During operation, the RAM heat exchanger may removeheat from any heat exchange fluid therein by flowing a certain amount ofRAM air over the RAM heat exchanger. Further, the bleed air heatexchanger is generally an “air to heat exchange fluid” heat exchangerwhich flows, e.g., bleed air from an LP compressor over heat exchangefluid to remove heat from the heat exchange fluid.

For the embodiment of FIG. 2, the one or more heat sink exchangers 108of the thermal management system 100 depicted includes a plurality ofindividual heat sink exchangers 108. More particularly, for theembodiment of FIG. 2, the one or more heat sink exchangers 108 includetwo heat sink exchangers 108 arranged in series. However, in otherexemplary embodiments, the one or more heat sink exchangers 108 mayinclude any other suitable number of heat sink exchangers 108. Forexample, in other exemplary embodiments, a single heat sink exchanger108 may be provided, at least three heat sink exchangers 108 may beprovided, at least four heat sink exchangers 108 may be provided, or atleast five heat sink exchangers 108 may be provided. Additionally, instill other exemplary embodiments, two or more of the heat sinkexchangers 108 may alternatively be arranged in parallel flow with oneanother.

Referring still to the exemplary embodiment depicted in FIG. 2, it willbe appreciated that the plurality of heat sink exchangers 108 and heatsource exchangers 106 are each selectively in thermal communication withthe thermal transport bus 102 (and selectively in fluid communicationwith the heat exchange fluid in the thermal transport bus 102). Moreparticularly, the thermal management system 100 depicted includes aplurality of bypass lines 110 for selectively bypassing each heat sinkexchanger 108 of the plurality of heat sink exchangers 108 and heatsource exchanger 106 of the plurality of heat source heat exchangers106. Each bypass line 110 extends between an upstream juncture 112 and adownstream juncture 114—the upstream juncture 112 located just upstreamof a respective heat sink exchanger 108 or heat source heat exchanger106, and the downstream juncture 114 located just downstream of therespective heat sink exchanger 108 or heat source heat exchanger 106.Additionally, each bypass line 110 meets at the respective upstreamjuncture 112 with the thermal transport bus 102 via a three-way heatsink valve 116. The three-way heat sink valves 116 each include an inletfluidly connected with the thermal transport bus 102, a first outletfluidly connected with the thermal transport bus 102, and a secondoutlet fluidly connected with the bypass line 110. The three-way heatsink valves 116 may each be a variable throughput three-way valve, suchthat the three-way heat sink valves 116 may vary a throughput from theinlet to the first and/or second outlets. For example, the three-wayheat sink valves 116 may be configured for providing anywhere betweenzero percent (0%) and one hundred percent (100%) of the heat exchangefluid from the inlet to the first outlet, and similarly, the three-wayheat sink valves 116 may be configured for providing anywhere betweenzero percent (0%) and one hundred percent (100%) of the heat exchangefluid from the inlet to the second outlet.

Notably, the three-way heat sink valves 116 may be in operablecommunication with a controller 115 of the turbofan engine 10 and/or ofan aircraft including the turbofan engine 10 through one or more wiredor wireless communications busses (depicted in phantom). The controller115 may bypass one or more of the one or more heat sink exchangers 108and/or heat source exchangers 106 based on, e.g., an operating conditionof the turbofan engine 10 and/or aircraft, a temperature of the heatexchange fluid, and/or any other suitable variables. Alternatively, thecontroller 115 may bypass one or more of the one or more heat sinkexchangers 108 and/or heat source exchangers 106 based on a user input.

Further, each bypass line 110 also meets at the respective downstreamjuncture 114 with the thermal transport bus 102. Between each heat sinkexchanger 108 and downstream juncture 114, the thermal transport bus 102includes a check valve 118 for ensuring a proper flow direction of theheat exchange fluid. More particularly, the check valve 118 prevents aflow of heat exchange fluid from the downstream juncture 114 towards therespective heat sink exchanger 108.

The thermal management system 100 of FIG. 2 may more efficiently removeheat from the various accessory systems of the turbofan engine 10 and/orthe aircraft, and in a more efficient manner. For example, in theexemplary embodiments including a plurality of heat sink exchangers 108having bypass capability, for example, the additional heat sinkexchangers 108 have the benefit of adding redundancy to the thermalmanagement system 100. For example, in the event of a failure of one ormore of the heat sink exchangers 108 or associated portions of thethermal transport bus 102, the heat exchange fluid may be routed aroundsuch failure and the system 100 may continue to provide at least someheat removal.

Referring now to FIG. 3, a close-up, cross-sectional view of a gasturbine engine including a thermal management system 100 in accordancewith an exemplary aspect of the present disclosure is provided. The gasturbine engine may be configured in a similar manner to the exemplaryturbofan engine 10 described above with reference to FIG. 1, andfurther, the thermal management system 100 may be configured in asimilar manner to the exemplary thermal management system 100 describedabove with reference to FIG. 2. Accordingly, the same or similar numbersmay refer to same or similar parts.

For example, as is depicted the exemplary gas turbine engine of FIG. 3generally includes a turbomachine 16 and an outer nacelle 50, with theturbomachine 16 at least partially surrounded by the outer nacelle 50.Moreover, the outer nacelle 50 defines a bypass airflow passage 56 withthe turbomachine 16 (i.e., between the outer nacelle 50 and theturbomachine 16), and more specifically, defines the bypass airflowpassage 56 between the outer nacelle 50 and an outer casing 18 of theturbomachine 16. Furthermore, the gas turbine engine includes an outletguide vane 52 extending between the outer nacelle 50 and theturbomachine 16, the outlet guide vane 52 supporting the turbomachine 16relative to the outer nacelle 50.

In such a manner, the gas turbine engine may be referred to as aturbofan engine (similar to the exemplary turbofan engine 10 of FIG. 1).Further, it will be appreciated from FIG. 3, and the discussion abovewith reference to FIG. 1, that the gas turbine engine may further definea relatively high bypass ratio, and therefore may be referred to as a“high-bypass” turbofan engine.

Referring still to FIG. 3, the exemplary turbomachine 16 depictedgenerally includes a compressor section, a combustion section 26, aturbine section, and an exhaust section 32. The compressor section,combustion section 26, turbine section, and exhaust section 32 togetherdefine at least in part a core air flowpath 37. Additionally, thecompressor section generally includes a high pressure (“HP”) compressor24, and the turbine section generally includes a low pressure (“LP”)turbine 30 and an HP turbine 28. The LP turbine 30 is coupled to, andconfigured to drive, an LP spool 36, and the HP turbine 28 is coupled toand configured to drive, an HP spool 34. Notably, the HP spool 34 isfurther coupled to the HP compressor 24, such that the HP turbine 28 maydrive the HP compressor 24 through the HP spool 34.

The turbomachine 16 further includes a fuel delivery system 120 forproviding a fuel flow to the combustion section 26 of the turbomachine16. For example, the exemplary fuel delivery system 120 generallyincludes one or more fuel nozzles 122 configured to provide a mixture offuel and air to a combustion chamber 124 of the combustion section 26,as well as a fuel pump 126 and a plurality of fuel lines 128. The fuelpump 126 may provide for the fuel flow through the plurality of fuellines 128 from a fuel source (not shown) to the plurality of fuelnozzles 122. Further, it will be appreciated that in at least certainexemplary embodiments, the fuel delivery system 120 may be used as aheat sink.

Moreover, as stated the exemplary gas turbine engine depicted includesthe thermal management system 100. In at least certain exemplaryembodiments, the thermal management system 100 of the gas turbine engineof FIG. 3 may be configured in a similar manner as the exemplary thermalmanagement system 100 described above with reference to FIG. 2. Forexample, the thermal management system 100 depicted generally includes aheat source heat exchanger 106 configured to collect heat from theturbomachine 16 during operation (i.e., collect heat from one or morecomponents of the turbomachine 16 during operation), a heat sink heatexchanger 108 configured to reject heat during operation, and a thermaltransport bus 102. The thermal transport bus 102 includes a heatexchange fluid configured to flow therethrough at a pressure within anoperational pressure range during operation (see, also, FIG. 2). Theheat source heat exchanger 106 and heat sink heat exchanger 108 are eachthermally coupled to the thermal transport bus 102, and morespecifically, to the heat exchange fluid within the thermal transportbus 102. In such a manner, the heat source heat exchanger 106 isoperable to transfer heat to the heat exchange fluid flowing through thethermal transport bus 102 and the heat sink heat exchanger 108 isconversely operable to transfer heat from the heat exchange fluidflowing through the thermal transport bus 102.

Notably, it will be appreciated that as used herein, the term “heatsource” and “heat sink” as used to describe a heat exchanger refer totypical operation of the heat exchanger with respect to the thermalmanagement system 100 and thermal bus 102. For example, a heat sourceheat exchanger 106 refers to a heat exchanger that is generally operableto provide heat to the thermal management system 100 and thermal bus102. However, relative to other systems to which it is thermallyconnected, the heat source heat exchanger 106 may act as a heat sink.Similarly, for example, a heat sink heat exchanger 108 refers to a heatexchanger that is generally operable to remove heat from the thermalmanagement system 100 and thermal bus 102. However, relative to othersystems to which it is thermally connected, the heat sink heat exchanger108 may act as a heat source. Further, during certain operations of theengine and thermal management system 100, the heat source heat exchanger106 may further be configured to act as a heat sink for the thermalmanagement system 100 and the heat sink heat exchanger 108 may furtherbe configured to act as a heat source for the thermal management system100.

More specifically, for the embodiment depicted, the heat source heatexchanger 106 is a first heat source heat exchanger 106A, and thethermal management system 100 further includes a second heat source heatexchanger 106B. The first heat source heat exchanger 106A is, for theembodiment shown, a waste heat recovery heat exchanger. Morespecifically, for the embodiment depicted, the waste heat recovery heatexchanger is positioned to be in thermal communication with the core airflowpath 37 within or downstream of, the turbine section, the exhaustsection 32, or both. More specifically, still, for the embodimentdepicted, the waste heat recovery heat exchanger is integrated into anaft strut/outlet guide vane 31 of the LP turbine 30 at a downstream endof the LP turbine 30 of the turbomachine 16 (i.e., downstream of allturbine rotor blades within the turbine section of the gas turbineengine). Accordingly, the waste heat recovery heat exchanger maygenerally capture heat from the flow of gases through an aft portion ofthe turbine section and/or from a flow of exhaust through the exhaustsection 32.

Further, it will be appreciated that the exemplary gas turbine enginedepicted includes a cooling air system 80 (see, also FIG. 1), and thesecond heat source heat exchanger 106B is configured as a cooling airsystem heat exchanger for the cooling air system. As will beappreciated, the cooling air system may generally provide an airflow 132from the compressor section to the turbine section, with such airflow132 being used as cooling air for the turbine section. The cooling airsystem heat exchanger may remove heat from the airflow 132 from thecompressor section prior to such airflow 132 being provided to theturbine section. In such a manner, the cooling air system heat exchangermay reduce a temperature of the airflow 132 being provided to theturbine section, such that the airflow 132 may be used more efficientlyas cooling air.

Also for the embodiment depicted, the heat sink heat exchanger 108 is afirst heat sink heat exchanger 108A, and the thermal management system100 further includes a second heat sink heat exchanger 108B. The firstheat sink heat exchanger 108A is, for the embodiment shown, a fuel heatexchanger thermally coupled to the fuel delivery system 120. Morespecifically, the fuel heat exchanger is thermally coupled to one of theplurality of fuel lines 128 of the fuel delivery system 120 such thatthe fuel heat exchanger may reject heat to a fuel flow therethrough. Itwill also be appreciated that although not depicted, the thermalmanagement system 100 may include one or more valves and lines allowingthe thermal management system 100 to bypass the fuel heat exchangerduring certain operating conditions (see FIG. 2 and discussion above).

Moreover, for the embodiment depicted, the second heat sink heatexchanger 108B is a bypass airflow heat exchanger integrated into, orcoupled to, one or more components positioned in, or otherwise exposedto, the bypass airflow passage 56. More specifically, for the embodimentdepicted, the bypass airflow heat exchanger is integrated into, orcoupled to, the outlet guide vane 52 of the gas turbine engine. In sucha manner, the bypass airflow heat exchanger may reject heat to theairflow through the bypass airflow passage 56 during operation.

Furthermore, as is also depicted schematically in FIG. 3, the thermalmanagement system 100 further includes a pump 104 and a heater 136. Thepump 104 is configured to increase a pressure of the heat exchange fluidwithin the thermal transport bus 102 such that the heat exchange fluidoperates as desired and/or flows through the thermal transport bus 102as desired. Additionally, the heater 136 is configured to increase atemperature of the heat exchange fluid flowing through the thermaltransport bus 102. For example, the heater 136 may be configured toincrease a temperature of the heat exchange fluid within the thermaltransport bus 102 during initial/startup operations, before the gasturbine engine has heated up. Such may therefore allow for the thermalengine system to operate in relatively low ambient temperatureconditions.

Referring now briefly to FIGS. 4 through 6, various embodiments of theexemplary heater 136 are depicted. For example, referring first to theexemplary embodiment of FIG. 4, a side, schematic view of a heater 136in accordance with an exemplary embodiment of the present disclosure, aswell as a section of the thermal transport bus 102 is provided. Asshown, in certain exemplary embodiments, the heater 136 may beconfigured as an electric resistance heater. The electric resistanceheater may be positioned in thermal communication with a length of thethermal transfer bus 102, either by physically contacting the thermaltransport bus 102, being positioned within the thermal transport bus102, or being positioned adjacent to thermal transport bus 102. In sucha manner, the electric resistance heater may heat the heat exchangefluid within the thermal transport bus 102.

Additionally, or alternatively, referring now to FIG. 5, providing aside, schematic view of a heater 136 in accordance with anotherexemplary embodiment of the present disclosure, the heater 136 mayutilize a heating fluid. For example, in certain embodiments, the heater136 may include a conduit 138 extending around and coaxially with alength of the thermal transport bus 102 (i.e., surrounding the thermaltransport bus 102, such that at least a length of the thermal transportbust 102 is positioned completely within the conduit 138 of the heater136). In such a manner, the heater 136 may flow a heating fluid throughthe conduit 138 and around the thermal transport bus 102 to heat thethermal transport bus 102 (and the heat exchange fluid therein). Morespecifically, the heating fluid within the conduit 138 of the heater 136may transfer heat to the thermal transport bus 102, which may in turntransfer heat to the exchange fluid within the thermal transfer bus.

Additionally, or alternatively, still, referring now to FIG. 6,providing a side, schematic view of a heater 136 in accordance with yetanother exemplary embodiment of the present disclosure, the heater 136may include a conduit 138 positioned within the thermal transport bus102 (i.e., positioned completely within the thermal transport bus 102,such the thermal transport bus 102 completely surrounds the conduit 138of the heater 134). For example, similar to the embodiment of FIG. 5,the heater 136 of FIG. 6 may utilize a heating fluid within the conduit138. The conduit 138, which may be positioned within the thermaltransport bus 102, extends generally coaxially with the thermaltransport bus 102, such that the conduit 138 is completely surrounded bythe thermal transport bus 102 and the heat exchange fluid therein. Theconduit 138 may carry the heating fluid, which may transfer heatdirectly to the heat exchange fluid within the thermal transport bus102. It will be appreciated, however, that in other exemplaryembodiments, any other suitable heater 136 may be utilized for heatingthe heat exchange fluid within the thermal transport bus 102, oralternatively, no heater 136 may be utilized.

Referring now back to FIG. 3, it should be appreciated that in otherexemplary embodiments of the present disclosure, the thermal managementsystem 100 may be configured in any other suitable manner. For example,in other embodiments, the thermal management system 100 may include anyother suitable number or type of heat source heat exchanger 106, as wellas any other suitable number or type of heat sink heat exchanger 108, asdiscussed above with reference to FIG. 2. Further, the thermalmanagement system 100 may also be integrated into, or utilized with, anyother suitable gas turbine engine.

Referring still to FIG. 3, it will be appreciated that in order tooperate in accordance with one or more of the exemplary aspectsdescribed herein, i.e., to accept heat from the one or more heat sourcesdescribed herein, the heat exchange fluid configured to flow through thethermal transport bus 102 must be capable of operating over a relativelylarge operating temperature. For example, the heat exchange fluid mustbe able to flow at relatively low temperatures, e.g., during startupoperations wherein the ambient temperature is relatively low, andfurther must not degrade at relatively high temperatures, e.g., duringhigh-power operating modes wherein a relatively large amount of, e.g.,waste heat is recovered and/or heat is removed from the cooling airsystem and/or other operating system(s). Accordingly, the thermalmanagement system 100 described herein defines an operationaltemperature range at a pressure within an operational pressure range forthe heat exchange fluid utilized within the exemplary thermal managementsystem 100. The operational temperature range includes a lowertemperature limit less or equal to than about zero degrees Fahrenheit ata pressure within the operational pressure range and an uppertemperature limit greater than or equal to about 1000 degrees Fahrenheitat a pressure within the operational pressure range. It will beappreciated, that as used herein, the term “operational temperaturerange” refers generally to a temperature range at a pressure within theoperational pressure range at which the heat exchange fluid may flowthrough the thermal transport bus 102 without substantially degradingitself or the thermal management system 100. Substantially degradingitself refers to the heat exchange fluid breaking down or changing intoother substances (e.g., a lubrication oil coking), and substantiallydegrading the thermal management system 100 refers to prematurelywearing out the system such that sustained or repeated use at theconditions is not possible.

For example, in at least certain exemplary embodiments, in order toallow for such a relatively large operating temperature range, the heatexchange fluid may be a liquid metal alloy, a molten salt, a siliconeoil, an ionic fluid, a pressurized gas, or a supercritical gas. Forexample, when the heat exchange fluid is a liquid metal alloy, the term“operational temperature range” may refer to a temperature range atwhich the liquid metal alloy remains in liquid form at a pressure withinthe operational pressure range, i.e., a temperature range higher than amelting point of the liquid metal alloy at a pressure within theoperational pressure range and lower than a boiling point of the liquidmetal alloy at a pressure within the operational pressure range. Forexample, the liquid metal alloy may be a eutectic alloy, such as aeutectic alloy including gallium, indium, tin, and/or other elements inany potential mass ratio. For example, the liquid metal alloy may be asubstantially gallium, indium, and tin eutectic alloy, such as theeutectic alloy known as GALINSTAN, available from Geratherm Medical AG.When the heat exchange fluid is a liquid metal alloy, the operationalpressure range may be relatively low, such as greater than zero poundsper square inch and less than about 500 pounds per square inch. Morespecifically, the operational pressure range may be a minimum pressurerange required to provide the desired flow rate through the thermaltransport bus 102.

Alternatively, as stated, in other exemplary embodiments, the exchangefluid may instead be a supercritical gas. For example, in certainexemplary embodiments, the supercritical gas may be a carbon dioxidegas. However, any other suitable supercritical gas may be used. When theheat exchange fluid is a supercritical gas, the term “operationtemperature range” may refer to a temperature range within which thesupercritical gas remains above its critical point at a pressure withinthe operational pressure range, and below a temperature limit for thethermal management system 100. For example, when a supercritical gas isutilized as a heat exchange fluid, the operational pressure range willbe relatively high. For example, it will be appreciated that thesupercritical gas will define a critical point pressure (i.e., apressure above which the fluid does not have distinct liquid and gasphases, provided it is above a critical point temperature). With such anexemplary aspect, the operational pressure range may be greater than thecritical point pressure of the supercritical gas, such as (depending onthe particular supercritical gas used) greater than about 700 pounds persquare inch and less than about 8000 pounds per square inch. Forexample, in certain exemplary aspects, the operational pressure rangemay be greater than about 1000 pounds per square inch, greater thanabout 1500 pounds per square inch, greater than about 2000 pounds persquare inch, or greater than about 2500 pounds per square inch and lessthan about 4,500 pounds per square inch.

Accordingly, it will be appreciated that the lower temperature limit ofthe operational temperature range defined by the heat exchange fluid ata pressure within the operational pressure range may be between about−100 degrees Fahrenheit and about −5 degrees Fahrenheit, such as lessthan about −10 degrees Fahrenheit, such as less than about −15 degreesFahrenheit, such as less than about −25 degrees Fahrenheit, such as lessthan about −50 degrees Fahrenheit. Furthermore, in at least certainexemplary embodiments, the upper temperature limit of the operationaltemperature range defined by the heat exchange fluid at a pressurewithin the operational pressure range may be between about 1000 degreesFahrenheit and about 1800 degrees Fahrenheit, such as greater than about1050 degrees Fahrenheit, such as greater than about 1100 degreesFahrenheit, such as greater than about a 1150 degrees Fahrenheit, suchas greater than about 1200 degrees Fahrenheit. Further, the uppertemperature limit of the operational temperature range may be less thanabout 3000 degrees Fahrenheit.

Inclusion of heat exchange fluid in accordance with one or more theseexemplary embodiments may allow for a thermal management system that iscapable of more efficiently managing thermal needs of the gas turbineengine.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A gas turbine engine comprising: a turbomachineincluding a compressor section, a combustion section, a turbine section,and an exhaust section arranged in serial flow order and togetherdefining a core air flowpath through the turbomachine; and a thermalmanagement system comprising a first heat exchanger integrated into acomponent of the turbine section, the component positioned inward from acasing in a radial direction and positioned within the core air flowpathof the turbine section of turbomachine, the first heat exchangerconfigured to collect heat from the turbomachine during certainoperations; a second heat exchanger; and a thermal transport bus havinga heat exchange fluid configured to flow through the thermal transportbus at a pressure within an operational pressure range, the first heatexchanger and the second heat exchanger each fluidly coupled to thethermal transport bus such that the first heat exchanger is operable totransfer heat to the heat exchange fluid during at least the certainoperations and the second heat exchanger is operable to transfer heatfrom the heat exchange fluid during at least certain operations, thethermal management system defining an operational temperature range forthe heat exchange fluid, the operational temperature range having alower temperature limit less than zero degrees Fahrenheit at a pressurewithin the operational pressure range and an upper temperature limit ofat least 1000 degrees Fahrenheit at a pressure within the operationalpressure range.
 2. The gas turbine engine of claim 1, wherein the firstheat exchanger is a heat recovery heat exchanger.
 3. The gas turbineengine of claim 2, wherein the heat recovery heat exchanger is a wasteheat recovery heat exchanger configured to capture heat from the coreair flowpath within or downstream of the turbine section, the exhaustsection, or both.
 4. The gas turbine engine of claim 2, wherein theturbomachine comprises a fuel delivery system, and wherein the secondheat exchanger is a fuel heat exchanger thermally coupled to the fueldelivery system.
 5. The gas turbine engine of claim 1, wherein theturbomachine comprises a cooling air system, and wherein the first heatexchanger is thermally coupled to the cooling air system.
 6. The gasturbine engine of claim 1, wherein the lower temperature limit of theoperational temperature range at a pressure within the operationalpressure range is between about −100 degrees Fahrenheit and about −5degrees Fahrenheit.
 7. The gas turbine engine of claim 6, wherein theupper temperature limit of the operational temperature range at apressure within the operational pressure range is between about 1000degrees Fahrenheit and about 1800 degrees Fahrenheit.
 8. The gas turbineengine of claim 1, wherein the heat exchange fluid is at least one of aliquid metal alloy, a molten salt, a silicone oil, an ionic fluid, apressurized gas, or a supercritical gas.
 9. The gas turbine engine ofclaim 1, wherein the heat exchange fluid is a liquid metal alloy, andwherein the operational pressure range is greater than zero pounds persquare inch and less than about 500 pounds per square inch.
 10. The gasturbine engine of claim 1, wherein the heat exchange fluid is a eutecticmetal alloy comprising gallium, indium, and tin.
 11. The gas turbineengine of claim 1, wherein the heat exchange fluid is a pressurized gas.12. The gas turbine engine of claim 1, wherein the heat exchange fluidis a supercritical gas, wherein the supercritical gas defines criticalpoint pressure, wherein the operational pressure range is greater thanthe critical point pressure of the supercritical gas and up to about8000 pounds per square inch.
 13. The gas turbine engine of claim 12,wherein the heat exchange fluid is a supercritical gas, and wherein theoperational pressure range is greater than about 1000 pounds per squareinch and less than about 8000 pounds per square inch.
 14. The gasturbine engine of claim 1, further comprising: an outer nacelle at leastpartially surrounding the turbomachine; and an outlet guide vaneextending between the outer nacelle and the turbomachine, wherein thesecond heat exchanger is integrated into, or coupled to, the outletguide vane.
 15. The gas turbine engine of claim 1, wherein the thermalmanagement system further comprises a heater thermally coupled to thethermal transfer bus for heating the thermal transfer fluid within thethermal transfer bus.
 16. The gas turbine engine of claim 1, wherein thefirst heat exchanger is a first heat source heat exchanger, wherein thethermal management system further comprises a second heat source heatexchanger, wherein the first heat source heat exchanger is a waste heatrecovery heat exchanger, and wherein the second heat source heatexchanger is a cooling air system heat exchanger.
 17. The gas turbineengine of claim 1, wherein the second heat exchanger is a fuel heatexchanger, and wherein the thermal management system further comprises abypass airflow passage heat exchanger.
 18. A gas turbine enginecomprising: a compressor section, a combustion section, a turbinesection, and an exhaust section arranged in serial flow order andtogether defining a core air flowpath; a fuel delivery system forproviding a flow of fuel to the combustion section; and a thermalmanagement system comprising a first heat exchanger integrated into acomponent of the turbine section, the component positioned inward from acasing in a radial direction and positioned within the core air flowpathof the turbine section, the first heat exchanger configured to collectheat from within or downstream of the turbine section, the exhaustsection, or both; a second heat exchanger in thermal communication withthe fuel delivery system for transferring heat to the flow of fuelprovided to the combustion section; and a thermal transport bus having aheat exchange fluid flowing therethrough, the first heat exchanger andsecond heat exchanger each fluidly coupled to the thermal transport bus.19. The gas turbine engine of claim 18, wherein the thermal managementsystem defines an operational temperature range for the heat exchangefluid, the operational temperature range having a lower temperaturelimit less than zero degrees Fahrenheit at a pressure within theoperational pressure range and an upper temperature limit of at least1000 degrees Fahrenheit at a pressure within the operational pressurerange.
 20. The gas turbine engine of claim 19, wherein the heat exchangefluid is at least one of a molten salt, a silicone oil, an ionic fluid,a liquid metal alloy, or a supercritical gas.